International Journal for Research in Engineering Application & Management (IJREAM) ISSN : 2454-9150 Vol-05, Issue-09, Dec 2019 Mathematical modeling and analysis of a

(LH2/LOX) combustion chamber with variable mass flow and wall temperature

ABHAY KUMAR, M. Tech Scholar Mechanical Engg. Dept. SRK , Bhopal, India. Dr. Manoj Chopra, Professor & HOD, RKDF Institute of Science & Technology, Bhopal, India. Dr. S.S. PAWAR, Professor RKDF Institute of Science & Technology, Bhopal, India.

ABSTRACT - The study of the subject on which the thesis is been written is for a cryogenic engine. Cryogenic is a branch of physics, which deals with the production and behaviour of materials at very low temperatures. Cryogenic being a vast branch of physic has many applications. The application with which or study is concerned cryogenic . Further explanation for cryogenic will be elaborated in this study. Our primary reason for doing this study is to assess the pressure and optimise the exact inlet pressure to face up to the combustion chamber for a constant 50KN thrust of and gravitational impact on the cryogenic combustion chamber. This study has been done by employing a superb technique that is commonly used currently a day’s known as CFD tools Fluent 14.5. For fluid flow system and chemical reaction model, Heat transfer and Mass transfer problem. Simulation results provide us acceptable contour results that describe the entire system in correct manner these results additionally validate the reference paper worth therefore our methodology is additionally validate that shows our case, giving boundary condition and geometrical parameters are correct. By CFD results we can conclude that micro-gravitational effect on cryogenic combustion chamber is negligible because velocity of such type of engine is very high as compare to increased velocity value due to lower gravity but increased value of pressure is definitely the positive results in this study. In current study we analyzed the CFD model of a cryogenic combustion chamber for various chamber pressures 40bar initially then changing the pressure by 50bar and 60 bar in Fluent software system. In current analysis we have a tendency to investigate the result of various microgravity through numerical investigation by victimisation gravity mechanism model in Fluent. Our main aim is to research the result of pressure variation on the chamber for a constant 50 KN thrust. Simulation results predict the flow development and turbulence model of a system. we have a tendency to see from the contour results of pressure, temperature and rate that once we increase the recess pressure from 40bar to 50bar and 60 bar the combustion chamber pressure is additionally will increase and contour rate also hyperbolic attributable to turbulence. During this whole method system is stable all the time this shows chamber is face up to or safe for constant 50KN thrust that is our main objective of the study.

Keywords: /, Combustion Device, Cryogenic Combustion, High Pressure, Diagnostic, Modelling, CFD, etc.

I. INTRODUCTION compared to other type of . are high energetic propellant combination, also the future space Payload capacity can be increased with the propulsion mission needs a propulsion system with the high thrust and system having higher , in general liquid reusable capacities, like Main Engine propellant engines results in longer burning time than (SSME), which can also leads to reduction in launch cost conventional solid rocket engines which result in higher per kg of payload with the highly improved payload specific impulse. Liquid propellants can be classified based capacity. As above said space shuttle main engine uses on the storable conditions, namely earth storable, cryogenic, and as a propellant space storable etc. Among these cryogenic propellants are combination with high specific impulse over the other widely used because of their high specific impulse cryogenic propellant combinations. Apart from the LOX/

162 | IJREAMV05I0957007 DOI : 10.35291/2454-9150.2019.0555 © 2019, IJREAM All Rights Reserved. International Journal for Research in Engineering Application & Management (IJREAM) ISSN : 2454-9150 Vol-05, Issue-09, Dec 2019

LH2 combination, LOX/ RP- 1 () semi- cryo II. LITERATURE REVIEW combination is widely used in Russian launch vehicle engines of RD- 107, RD- 108and RD- 461 in the Senthilkumar et al. (2013)[1] The basic concept of a rocket launch vehicles, which offers slightly lower specific engine relays on the release of the internal energy of the impulse than LOX/ LH2 combination with advantage of propellant molecules in the combustion chamber, the handling. The demand of increasing payloads and acceleration of the reaction product and finally the release increasing size of satellite increases the need of more of the hot gases at the highest possible velocity in the efficient rocket propulsion system. The cost of the payload convergent/divergent nozzle. Liquid rocket engines burn can be decreased by implementing new reusable launching propellants, which undergo chemical reactions to convert systems with the intentions to increase the efficiency of the the stored chemical energy to thermal energy which results launching system. For such condition high energetic in the generation of thrust. Thrust chamber of cryogenic cryogenic propellant combination provides suitable thrust engine is modelled at a chamber pressure of 40 bar and to weight ratio with high specific impulse than current thrust of 50KN to reduce the high temperature and pressure existed earth storable propellants. Reduction in launch cost in the combustion chamber. CFD analysis is done to show per kg of payload with the highly improved payload the pressure and temperature variation in the thrust chamber capacity. As above said space shuttle main engine uses modelled for 50KN thrust and chamber pressure of 40 bar. Liquid oxygen and liquid hydrogen as a propellant The design of rocket engine should be such that it should combination with high specific impulse over the other withstand the high pressure and high temperature of the cryogenic propellant combinations. Apart from the LOX/ combustion chamber. Ashwini, and Prabhakaran (2015) [2] LH2 combination, LOX/ RP- 1 (Kerosene) semi- cryo This paper highlights about the rocket engine involving the combination is widely used in Russian launch vehicle use of cryogenic technology at a cryogenic temperature engines of RD- 107, RD- 108and RD- 461 in the Soyuz (123 K). This basically uses the liquid oxygen and liquid launch vehicles, which offers slightly lower specific hydrogen as an oxidizer and fuel, which are very clean and impulse than LOX/ LH2 combination with advantage of non-pollutant fuels compared to other fuels handling. The demand of increasing payloads and like Petrol, Diesel, , LPG, CNG, etc., sometimes, increasing size of satellite increases the need of more liquid nitrogen is also used as fuel. The efficiency of the efficient rocket propulsion system. The cost of the payload rocket engine is more than the jet engine. As per the can be decreased by implementing new reusable launching Newton’s third law of mechanics, the thrust produced in systems with the intentions to increase the efficiency of the rocket engine is outwards whereas that produced in jet launching system. For such condition high energetic engine is inwards. Poschner and Pfitzner (2009) [3] M cryogenic propellant combination provides suitable thrust Poschner and M Pfitzner in this paper studied real gas as to weight ratio with high specific impulse than current . Due to the high pressures and very low existed earth storable propellants. The propellant injection temperatures of the propellants in modern rocket combination of liquid hydrogen and liquid oxygen has high combustors real gas effects play an important role in rocket specific impulse compared to other propellant combinations combustion simulation. These have to be accurately have been used in many rocket engines. The thermal modelled in combustion CFD simulations to enable an analysis is a major issue in at the channel exit. Those accurate prediction of the performance of the rocket parameters are important for the design of injectors and of combustion chamber. There research was based on the coolant pump. The design of a liquid rocket engine, theoretical model. They started with a simple experimental because the prediction of peak heat-flux from the model and conducted simulation on cfx. The gases they combustion gases to the engine wall is necessary to ensure used were liquid oxygen at 85 K and gaseous hydrogen at the structural integrity of the combustion chamber. The 257 K. They achieved a pressure of about 100 bar in need for thermal analysis is essentially important to extend combustion chamber at the time of combustion. Foe the engine life by effective and efficient cooling system. injecting gases they used co axial injector. Vivek Gautam Moreover, the analysis of the cooling channel flow is (2007) [4] The study submitted by Vivek Gautam (2007) essential to predict not only the efficiency of the coolant, worked profoundly on rocket jet propulsion two phase co but also the coolant temperature and pressure. The design of axial injector. His study marked a significant point on thrust chamber consists of many parameters and detail combustion and injection of propellant in combustion calculations, using basic geometric parameters are adequate chamber. The previous studies that were conducted lack to understand the regenerative cooling effect of the system. informative data. On such type of flows is either incomplete For the built-up of gas dynamic profile of the combustion or ambiguous. Previous models available on the same study chamber, it is necessary to give some input data to the lacked operating conditions. Operating conditions were system such as thrust (at sea level), chamber pressure, very narrow and limited. This part of this literature review ambient pressure and propellant components. is focused on the research performed on non-reacting two phase coaxial jets with cryogenic fluids such as liquid oxygen (LOX) and liquid nitrogen (LN2). These

163 | IJREAMV05I0957007 DOI : 10.35291/2454-9150.2019.0555 © 2019, IJREAM All Rights Reserved. International Journal for Research in Engineering Application & Management (IJREAM) ISSN : 2454-9150 Vol-05, Issue-09, Dec 2019 experiments were performed to analyze the flow, different microgravity through numerical investigation by atomization and mixing characteristics of shear coaxial using gravity mechanism model in Fluent. Our main aim is injectors typically used in cryogenic rocket engines. Vivek to investigate the effect of pressure variation on the Gautam (2007) on studying Bellan (2000) [7] found that chamber for a constant 50 KN thrust. CFD analysis will be she performed an extensive review on subcritical and done to show the pressure, temperature and velocity supercritical fluid atomization behaviour and modelling variation in the thrust chamber modeled for 50KN thrust issues. She showed that fluid jet disintegration under and chamber pressure of 40, 50 and 60 bar. The design of supercritical conditions is fundamentally different from the rocket engine should be such that it should withstand the much studied subcritical liquid atomization. Instead of the high pressure and high temperature of the combustion subcritical wave formation at the surface of the liquid chamber. resulting from the relative velocity between the liquid and ESULTS NALYSIS gas, ensuing in the subsequent instability and the further IV. R A breaking of the liquid sheet, under supercritical conditions CASE – I => Pressure – 40 Bar, Temp- 3420 K, g= 9.81 the fluid disintegrates in a remarkably different fashion. The optical data shows wispy threads of fluid emanating from the jet boundary and dissolving into the surrounding fluid. The existing information on two phase flows and atomization behaviour is of qualitative nature for supercritical fluids and considerable experimental work is necessary to provide data appropriate for model validation. She also discussed and evaluated the accuracy of existing models based on specific aspects supercritical fluids, such as intrinsic transient behaviour, lack of a material surface, real gas equations of state, mixture non-idealistic, increased solubility, Soret and Dufour effects and high pressure transport properties. Another researcher Woodward (1994)[8] used X-ray radiography and instantaneous Figure 1 Pressure at 40 Bar Inlet Pressure @ g=9.81 imaging technique to measure the potential core length of LOX stimulant (KI in aqueous solution) for a large range of Reynolds number, Weber number and density ratios. His measurements included two different approaches to measure the intact core length of LOX stimulant. The first technique was thresh-holding of jet images to reveal the core images corresponding specific liquid integrated thickness while the second technique was de convolution of time averaged images to obtain mean liquid volume fraction distributions. Although both the techniques had inherent unknown errors, they provided some very important measurements of cryogenic flow stimulants for the first time. Figure 2 Velocity at 40 Bar Inlet Pressure @ g=9.81 III. OBJECTIVE CASE – II => Pressure – 50 Bar, Temp- 3420 K, g= 9.81 Due to their high specific impulse, Oxygen/Hydrogen systems are often used to increase launcher performance and will probably continue to be the preferred option for the decades. In order to get insight into complex processes involved in the combustion of such propellants, research efforts have been conducted in both France and Germany including theoretical and experimental activities. This article deals with recent progress achieved by research teams from both countries in high pressure Oxygen/Hydrogen (O2/H2) combustion systems. In current study we will analyze the CFD model of a cryogenic combustion chamber for different chamber pressure 40bar initially and then changing the pressure by 50bar and 60 bar Figure 3 Pressure at 50 Bar Inlet Pressure @ g=9.81 in Fluent software. We will also investigate the effect of

164 | IJREAMV05I0957007 DOI : 10.35291/2454-9150.2019.0555 © 2019, IJREAM All Rights Reserved. International Journal for Research in Engineering Application & Management (IJREAM) ISSN : 2454-9150 Vol-05, Issue-09, Dec 2019

CASE – V => Pressure – 60 Bar, Temp- 3420 K, g1= 0.5g

Figure 4 Velocity at 50 Bar Inlet Pressure @ g=9.81 Figure 9 Pressure at 60 Bar Inlet Pressure @ g1=0.5g CASE – III => Pressure – 60 Bar, Temp- 3420 K, g= 9.81

Figure 10 Velocity at 60 Bar Inlet Pressure @ g1=0.5g CASE – VI => Pressure – 60 Bar, Temp- 3420 K, g = Figure 5 Pressure at 60 Bar Inlet Pressure @ g=9.81 1 0.25g

Figure 6 Velocity at 60 Bar Inlet Pressure @ g=9.81 Figure 11 Pressure at 60 Bar Inlet Pressure @ g1=0.25g CASE – IV => Pressure – 60 Bar, Temp- 3420 K, g1= 0.75g

Figure 12 Velocity at 60 Bar Inlet Pressure @ g1=0.25g

Figure 7 Pressure at 60 Bar Inlet Pressure @ g1=0.75g V. CONCLUSION In current study we will analyze the CFD model of a cryogenic combustion chamber for different chamber pressure 40bar initially and then changing the pressure by 50bar and 60 bar in Fluent software. In current research we investigate the effect of different microgravity through numerical investigation by using gravity mechanism model in Fluent. Our main aim is to investigate the effect of pressure variation on the chamber for a constant 50 KN thrust. Figure 8 Velocity at 60 Bar Inlet Pressure @ g1=0.75g

165 | IJREAMV05I0957007 DOI : 10.35291/2454-9150.2019.0555 © 2019, IJREAM All Rights Reserved. International Journal for Research in Engineering Application & Management (IJREAM) ISSN : 2454-9150 Vol-05, Issue-09, Dec 2019

Simulation results predict the flow phenomenon and Study, Characteristics on the Surface of Round Liquid Jet”, turbulence model of a system. We see from the contour Experiments in Fluids Vol. 16, 1994. results of pressure, temperature and velocity that when we [12] Branam, R. and Mayer, W.: “Length Scales in increase the inlet pressure from 40bar to 50bar and 60 bar Cryogenic Injection at Supercritical Pressure”, Experiments the combustion chamber pressure is also increases and in Fluids Vol. 33, 2002, streamline velocity also increased due to turbulence. In this whole process system is stable all the time this shows [13] Porcheron, E., Carreau, J. L., Prevost, L., Visage, D. chamber is withstand or safe for constant 50KN thrust Le, Roger, F.,: “Effect of Injection Gas Density on Coaxial which is our main objective of the study. Liquid Jet Atomization”, Atomization and Sprays, Vol. 12, Issue 1, 2002 REFERENCES [14] Chehroudi, B., Cohn, R., and Talley, D. G.: [1] S. Senthilkumar , Dr. P. Maniiarasan ,Christy Oomman “Cryogenic Shear Layers: Experiments and Initial Growth Jacob ,T. Vinitha Design And Analysis Of Thrust Chamber Rates of Round Cryogenic Jets at Subcritical and Of A 2013 Supercritical Pressures”, International Journal of Heat and [2] A.m. Ashwini, M. Prabhakaran “Cryogenic technology Fluid Flow, Vol. 23, No. 5, 2002 in rocket engines” B.E. Final year, Dept of Aeronautical [15] Candel, S., Juniper, M., Singla, G., Scouflaire, P. and Engineering, Dr. Pauls Engineering College. 2015 Rolon, C.: “Structure and Dynamics of Cryogenic Flames [3] M. Poschner, M. Pfitzner “CFD-Simulation of at Supercritical Pressure”, Combustion Science and supercritical LOX/GH2 combustion considering consistent Technology, 178, 2006 real gas thermodynamics” Thermodynamics Institute, [13] Habiballah, M., Orain, M., Grisch, F., Vingert, L., Faculty for Aerospace Engineering, University of the Gicquel, P.: “Experimental Studies of High-pressure Federal German Armed Forces , Munich 2009 Cryogenic Flames on the Mascotte Facility”, Combustion [4] Vivek Gautam “Flow And Atomization Characteristics Science and Technology, 178, 2006 Of Fluid From A Co Axial Rocket Injector” Graduate [17] Ivancic, B. and Mayer, W.: “Time and Length Scales School of the University of Maryland, College Park,2007 of Combustion in Liquid Rocket Thrust Chambers”, [5] Oschwald, M., Smith, J. J., Branam, R., Hussong, J., Journal of Propulsion and Power, Vol. 18, No. 2, March- Schik, A., Chehroudi, B. and Talley, D.: “Injection of April. 2002, Fluids into Supercritical Environments”, Combustion [18] Oefelein, J. C. and Yang, V.: “Modeling High Pressure Science and Technology, Vol. 178, 2000 Mixing and Combustion Processes in Liquid Rocket [6] Bellan, J.: “Supercritical (and subcritical) Fluid Engines”, Journal of Propulsion and Power, Vol. 14, No. 5, Behavior and Modeling: Drops, Streams, Shear and Mixing September-October 1998 Layers, Jets and Sprays”, Progress in Energy and Combustion Science, Vol. 26, 2000 [7] Woodward, R.: “Doctoral Dissertation”, Submitted to Department of Mechanical Engineering, Penn State University, 1994 [8] Glogowski, M., Bar-Gill, M., Puissant, C., Kaltz, T., Milicic, M., and Micci, M., “Shear Coaxial Injector Instability Mechanisms,” 30th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Indianapolis, IN, June 27-29, 1994, AIAA-94-2774 [9] Vingert, L., Gicquel, P., Lourme, D., and Menoret, L.: “Coaxial Injector Atomization”, Progress in Aeronautics and Astronautics, Vol. 169, AIAA, 1994. [10] Carreau, J. L., Porcheron, Prevost, L., E., Visage, D. Le and Roger, F., Gicquel, P. and Vingert, L.:” Atomization of a Liquid Oxygen Jet by a Coaxial Inert Gas Jet”, 14th Conference on Liquid Atomization and Spray Systems, Florence, 1997 [11] Mayer, W.: “Coaxial Atomization of a Round Liquid Jet in a High Speed Gas Stream: A Phenomenological

166 | IJREAMV05I0957007 DOI : 10.35291/2454-9150.2019.0555 © 2019, IJREAM All Rights Reserved.